Missile guidance method and apparatus



Nov. 25, 3969 J. CHU

MISSILE GUIDANCE METHOD AND APPARATUS .5 Sheets-Sheet 1 Filed July El,1954 w mxk a/WH ATTORNE S huomqh Q/fi 222 33 6 .528 3.5.35. N .w\.&

Nov. 25, 1969 J. CHU 3,480,233

MISSILE GUIDANCE METHOD AND APPARATUS Filed July 21,, 1954 .5Sheets-Sheet 2 i INVENTOR LAN J. mu k flag/ BY Q I? W ATTORNEYS Nov. 25,1969 3,480,233

Filed July 21, 1954 5 Sheets-Sheet 5 INVENTOR L a. 6H0" ATTORNEYS NOV.25, 1969 J c u 3,480,233

MISSILE GUIDANCE METHOD AND APPARATUS Filed July 21, 1954 5 Sheets-Sheet4 ATTORNEY Nov. 25, 1969 J. CHU 3,480,233

MISSILE GUIDANCE METHOD AND APPARATUS Filed July 2]., 1954 .5Sheets-Sheet 5 Om 0 JACK 522.53 @515 22 m2: Y B

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ATTORNEY? United States Patent 3,480,233 MISSILE GUIDANCE METHOD ANDAPPARATU Lan J. Chu, Lexington, Mass., assignor, by mesne assign--ments, to the United States of America as represented by the Secretaryof the Navy Filed July 21, 1954, Ser. No. 444,931 Int. Cl. F41g 7/.00;F42b 15/02; G06f 15/50 US. Cl. 244-318 6 Claims ABSTRACT OF THEDISCLOSURE The disclosure is directed to a missile guidance systemincorporating radiant energy detection devices, which feed signalprocessing circuitry for phase differentiating of the signals in amanner to provide steering signals for application to the controlsurfaces of the missile for course changing to effect continuous pursuitof a target by the missile.

This invention relates to the guidance of missiles, and particularly tothe problem of directing and maintaining a missile on a collision coursewith respect to a target traveling through space, which target possessesthe capability of maneuvering to avoid collision with such a missile.

An object of the invention is to provide a novel method of guiding amissile toward a course being followed by a target maneuvering throughspace, the method disclosed being so devised as to cause the missile tocontinually change its course to conform to each successive deviation inthe targets course.

Another object is to provide a guidable missile of novel construction,including steering apparatus responsive to phase differentiated radiantenergy signals received from a detected target, and operable to changethe course of the missile continually so long as the detected targetcontinues evasive maneuvers.

A third object is to provide, as part of the novel guidance method abovereferred to, the step of applying lateral acceleration to a missile inresponse to variations in the rate of change in the significantcharacteristics of intelligence reflected from a detected target, theresulting lateral acceleration serving as a means to direct the missileto a collision course, that is, a course most likely to bring aboutcollision between the missile and the detected target, on the basis ofthe intelligence being concurrently received by the missile concerningthe behavior of the target. i v

..A fourth object is to provide, as part of the novel guidance method,the step of converting intercepted radiant energy into missile guidinginformation, by processes including the utilization of fixed antennaelements carried by the missile, which processes eliminate the need ofactually computing the speed or relative direction of either the missileor the target.

A fifth object is to provide, as a specific part of the processes justreferred to, the step of varying the lateral acceleration of the missilein accordance with the rate of change of phase difference between signalenergy intercepted at opposite extremities of the fixed antenna arrayabove referred to; said variation in lateral acceleration being broughtabout by directing said signal energy through a seeker mechanism adaptedto convert said phase difference rate of change into electricalenergizing impulses of appropriate polarity and magnitude to actuate aservo mechanism in the proper direction, and to the proper degreerequired to cause the desired shift of externally mounted vanes, orsteering surfaces, about the pitch and yaw axes, respectively, of themissile control system.

A sixth object is to provide a novel combination and relationship ofactuating units for putting into practicethe novel missile guidancemethod herein disclosed.

Other objects and many of the attendant advantages of this inventionwill be readily appreciated as the same becomes better understood byreference to the following detailed description when considered inconnection with the accompanying drawings wherein:

FIG. 1 is a diagram indicating the circular course into which a missileis guided during the first stage of the novel method of operation hereindisclosed;

FIG. 2 is a diagram illustrating both the first and the second stages inthe missiles course to the target;

FIG. 3 is a diagram to assist in identifying certain symbols used in thedescription and claims;

FIG. 4 is a diagram illustrating the converging courses the missile andtarget may follow;

FIGS. 5, 5A, and 5B are diagrams illustrating the antenna and targetrelationships;

FIG. 6 is a perspective view of a missile embodying, and suitable forpractice of, the invention;

FIG. 7 is a schematic representation of actuating and control componentsassociated with one of the three control servos (namely, the yaw controlservo) which operate conjointly to execute the method of control hereindisclosed; and

FIG. 8 is a sectional view of the missile, showing the relativelocations of the major components.

By way of introduction to the present invention, it should be pointedout that the. performance of an unattended missile will necessarily belimited by the capacities inherent in the physical components carried bythe missile for target-seeking, propelling and guiding purposes. Thenovel method of operation herein proposed is believed to be Well adaptedto such inherent limitations, in that it can be executed with a minimumof reliance upon factors of variable or unpredictable performancecharacteristics. In fact, the method can be carried out by utilizationof simple steering mechanism controlled by reflected radar pulses, yetwithout the complications of rotating scanning antennas.

t of collision with the target. In other words, when w arrives at avalue equal to dfi/dt (hereinafter written as ,5), the missile is atthat instant on a collision course with the target, 5 representing theangle between the heading of the target and the line joining the targetand the missile I and 1 representing time. Hence only the derivative ofthe missile angle and not the angle itself, need be supplied to theradar seeker circuits. This discovery makes possible the elimination ofa scanning type of antenna. The steps leading to this mathematicalconclusion are developed in the following analysis of the problem, whichanalysis extion capacity being necessarily restricted by the mechanicaldesign of the missile. The same is true of the other limiting quantity,namely, the missiles velocity; but the linear velocity of the missile isnot readily controllable by automatic means, and, as above noted, themathematical conclusions herein expressed make possible the developmentof a control system operating through a single variable exclusively,namely, the lateral acceleration of the missile.

The following analysis of the problem will further explain how it ispossible to direct the missile to the calculated point of collision byvarying only a single characteristic of the missile, namely, its abilityto accelerate laterally.

Let V be the linear speed of the missile, and Am the maximum lateralacceleration that can be applied to the missile, (as by application ofturning effort to the external control surfaces). If this maximumacceleration is applied at the point of detection, and maintainedthereafter, the missile will follow a circular course of minimum radiusR, where the course proceeding either to the right or left, as shown inFIG. 1; the direction of turn depending upon whether the target isproceeding leftward or rightward from the missile-target axis.

In a three-dimensional view of the action, either circle shown in FIG. 1will generate a tore (doughnut) sweeping out a space whose width istwice the radius R, the axis of the tore being continuously shifted tomaintain tangency with the shifting course of the missile, until thetime is reached when the shifting axis points directly to the predictedpoint of collision (see FIG. 2). At that instant the missiles controlsystem should discontinue turning force application, and shouldthereafter apply a constant bearing effort to maintain the missile onthe collision axis.

Thus, the ideal course, as illustrated in FIG. 2 by the heavy line, hastwo stages, the first stage being a path of maximum curvature, and thesecond stage being a straight path tangential to the toroidal axis ofthe first path at that point thereon which represents the missilesposition at the instant of time when the missiles angular velocity hasincreased sufiiciently to equal its time rate of change in heading, withrespect to the missile-target axis; that is, when This course is theideal one because it is the shortest possible course and (assumingconstant missile velocity) places the point. of collision on thetoroidal axis of the turning missile at the earliest possible moment andhence at the greatest possible range, thereby assuring maximum strikingprobability. And as the second stage is of maximum length, the lateralacceleration requirements are minimized, so that less power need beaccumulated for application to the control servos.

The control system that will steer the missile along the ideal coursejust described will now be considered. The following factors areinvolved:

V=the speed of the missile,

v=the speed of the target,

B=the angle between the heading of the missile and the line joining thetarget and the missile,

a=the angle between the heading of the target and the line joining thetarget and the missile,

D=the distance between the target and the missile,

A=the lateral acceleration of the missile, and

w=the angular velocity of the missile (=A/ V) The following relationshipobtains between these quantities:

' DMJ sing D D where [3 is the time rate of change of B. Thefirst termofthe right-hand side is the rate of change of heading with respect tospace. The second term is that caused by the linear motion of themissile with respect to the target. The third term is that due tothelinear motion of the target with respect to the missile. The sum of thethree gives the total rate of the change of heading with respect to theline joining the missile and the target. The angular velocity of thetarget comes into the picture as a secondorder efiFect only.

Equation 2 can be rearranged as follows:

as v and a, which are not under the direct control of the missile. If,however, it is set equal to zero,

it is the equation for a constant-bearing collision course. Each term isthe common altitude of a triangle with angles equal to a and f3 andsides proportional to v and V, respectively. This is exactly what isneeded for the second stage of the ideal missile course. However, as theproper value of ,8 depends not only upon the velocity ratio but alsoupon the target angle cc, the proper value cannot be determined bydirect measurement.

When the right-hand side of Eq. 3 is not equal to zero, it gives ameasure of the deviation from a collision course. It follows that is therate of change of bearing with respect to a fixed line in the plane ofmaneuver.

Let it be assumed that the missile carries gyro mechanisms and radardetection means capable of measuring it gives a measure of hte deviationfrom a collision course;

respectively. At the instant the radar in the missile detects thetarget, and the quantities wand B are measured, the deviation from thecollision course can immediately be determined qualitatively. If (5)ispositive,

V sin fl v sin a and B should be increased so as to equalize the twoterms.

" In otherwo'rds, a positive a is needed. This can e imcomplished'byincreasing was shown in Eq. 2. As mentioned hereinbefore, it is desiredto have maximum 'ac-' celeration during the first stage of the approach.If the missile is traveling along a straight line initially will change,after. the application of the acceleration, from a, negative to apositive quantity. This maximum acceleration is to be maintained until(5) vanishes, at which instant the missile is heading along thecollision course. Then i H Since a: is not now equal to Zero, themissile will overshoot towards the other side of the collision course,and expression (5)" will become negative. A force is 'thenappliedthrough the servo to reverse the acceleration. With such a system, themissile will eventually oscillate about the collision course. The periodof the oscillation is'determined by the time lag present in transferringthe information regarding B from the radar system to the controlmechanism, and the time it takes to reverse the acceleration. Theshorter the period of oscillation the less the maximum deviation fromthe collision course.

If the target is maneuvering, the operation is not affected during thefirst stage of the approach. The missile is kept on the maximum lateralacceleration course until a collision course is reached based upon thevalues of a and 5 at that instant. After that, the lateral accelerationrequired for the missile is equal to the lateral acceleration of thetarget multiplied by a factor cos a/cos ,3. The missile will oscillateabout a course such that the average acceleration is equal to thatrequired to compensate for the maneuvering of the target.

As a numerical example, reference is made to FIG. 4 in which the targetis assumed to travel along a straight line.

Target velocity-1000 ft./sec.

Missile velocity2000 ft./ sec.

Maximum missile angular velocity5/sec.=87 mil./sec.

t the time during which the missile travels along a circular course tthe time during which the missile travels along a straight-line coursethe angle between the target and missile heading at the point ofcollision The absolute value of l B I will be about 2/sec, assuming avalue of 1st of about 3/ sec, for the period before the missile gets onthe collision course.

For t O, w 5

remains fairly constant and then approaches zero rapidly when themissile is near the collision course.

In order to reduce the amount of overshoot, the lateral accelerationmust be reduced when the missile nears the collision course. This may bedone by making the instantaneous angular velocity to proportional to01-6. The ratio of w to 04-3 must be several times unity for a fastmoving target. When t B is large, the angular velocity of the missile islimited by the maximum available acceleration. When the missile is nearthe constant-bearing course, the angular velocity is reducedproportionally with -B The amount of overshoot can also be reduced byapplying an acceleration that is positive or negative depending upon thesign of instead of that of where k is a positive constant less thanunity. When the missile is on the circular course, 3 has the same 'signas at but is numerically less than no. As the missile approaches theconstant-bearing course, 18 approaches 1 as a limit. Therefore, if k isless than unity Y willgo through zero sooner than will The accelerationstarts to change just before the missile reaches the collision course.This scheme, however, does not prevent the oscillation of the missileabout the collision course, but it does reduce th eperiod and eventuallydamp out the oscillation. The proper choice of the constant k dependsupon the time lag in the radar and servo system, and the time necessaryto reverse the acceleration.

From the preceding discussion, it is obvious that the only informationrequired of the radar is the rate of change of the angle 3, that is, thelead angle between the missile heading and the line of sight to thetarget. Since the angle information is not essential for the pursuit ofthe target, an antenna system that is fixed with respect to the missilemay be used, with a resultant simplification of the antenna and mountproblem, for either a ramjet or rocket type of missile.

As shown in FIGS. 5, 5A and 6, four identical antennas are mountedsymmetrically on the sides of the missile. The separation distancebetween each pair is approximately equal to the missile diameter plusthe length of diametrically opposed wings, which distance is smallcompared to the distance from the missile to the target. When a signalfrom the target arrivesat the missile, there is a phase difference of(21rd/ sin [3 between the two received signals at the antennas. As theangle 3 changes, the phase difference varies at a rate In other words,the two received signals will differ by a which is proportional to 3.This frequency difference, which is a measure of the time rate of changein the angle 8, determines the time interval during which the missileshould be maintained in the circular path constituting the first stageof maneuvering, as heretofore explained. Therefore, the two out-of-phasesignals received at vertical plane antenna elements A and B (andsimilarly as to those received at horizontal plane antenna elements Aand B need only be fed into suitable detection and phase timingcircuits, superimposed upon a suitable carrier wave, and then frequencymodulated, combined, and the resultant voltage selectively applied, atthe selected frequency, and with proper amplification, to operate apitch control servo unit (or yaw control servo unit, as the case maybe). One such arrangement is illustrated in FIG. 7 and controls theswing of rudder surfaces 1010 hingedly mounted on the vertical finscarrying the antenna elements A and B,,. In the embodiment heredescribed there are three such servo units, one for yaw, one for pitch,and one for roll control, that is, one servo for each of the threereference axes, X, Y and Z of FIGS. 5B and 7. The servo unit 11 forcontrol of yaw, only, is shown in FIG. 7. The other two, not shown, arepreferably of the same design, with each including a fluid pressureactuated piston or ram 11 to which hydraulic fluid is supplied from anaccumulator (not shown) at a rate, and in a direction, determined by themovement of the control valve 13 operated by a crossbar 14 swingingabout the axis of the shaft of a torque motor 15 to which is deliveredenergizing voltage whose direction and magnitude is controlled bypush-pull amplicos B) B fiers in an electronic unit represented by block16 in intermediate frequency by a common locol oscillator and amplifierto another input of said phase detectors 31 and 32 through a pair ofR.C. fixed phase shifters 33 and 34. The respective outputs of phasedetectors 31 and 32 will have a phase relation dependant on theinstantaneous phase relation of the signals from antennae A and B Saidoutputs are converted to a frequency modulated control signaled byapplying each to a balanced modulator, 37 and 38, and injecting a 2000cycle signal in fixed phase relation to each modulator. The phase of the2000 cycle signal at each modulator is determined by R.C. circuits 35and 36. The modulator outputs, when combined in block 17, produce acontrol signal having 2000 cycle carrier frequency modulated inaccordance with the rate of phase variation of the signals at theantannae. Amplitude modulation is removed in limiter 39 and themagnitude and sense of the phase rate is detected by discriminator 40.The discriminator output is compared with the yaw gyro output (discussedbelow) and delivers a control signal to torque motor 15 in response todifferences between the phase rate signal and the gyro signal.

An integrating gyroscope 46 (FIG. 7) whose precession is a measure ofangular deviations of the missile airframe from the prescribed referenceheading, is mouned in a cylindrical case (not shown) which is carriedhorizontally, with its axis on the X-axis of the missile, and iselectrically driven to rotate about the Y-axis of the missile. Asangular velocity is applied to the missile a precession torque isdeveloped about the X-axis, thereby rotating generator 48 and causing itto produce a voltage that is applicable to amplifier 16 by way ofconductor 49. This input via conductor 49 to amplifier 16, when added atthe amplifier to the command signal entering amplifier 16 from theseeker output circuit 51, results in the generation of quantity thenecessary servo command required for maintenance of the angular velocitydictated by the value set up by the phase comparison seeker circuits 21,22. The gyro unit includes a second generator 50 rotatable on the X-axisby feedback energy from circuit 52, which circuit 52 includes anamplifier 53 triggered by a pick-off unit 54 coupled to ram 11, tointroduce a rate correcting follow-up to the signal generating action ofgenerator 48. A similar gyro combination will perform a correspondingcompensating function with respect to the command signal voltagesupplied to the pitch servo unit by the seeker circuit associated withantennae A and B which seeker circuit will duplicate the one showndiagrammatically in FIG. 7 as associated with antennae A and B A thirdgyro combination of similar design will introduce the third (roll axis)compensating function by supplying operating signal voltage, derivedfrom the angular velocity of roll, to the roll servo unit controllingthe angular deflection of the ailerons 10b shown in FIG. 6.

FIG. B shows the geometrical relationships between the three missileaxes, X Y and Z and the lead angle L, which is the resultant of the twolead angle components L and L corresponding to the [3 and ,8 anglesshown in FIGS. 5 and 5A, respectively.

The general configuration of themissile, as shown in FIGS. 6 and 8, isnecessarily determined to a large extent by the requirements of theguidance system as hereinabove analyzed. Thus, the cruciform canardarrangement close to the forward nose, serving as a rigid mount for thefour homing antennae, is the most practical structure for carrying theyaw and pitch control surfaces 'and 10a, respectively, and thereforedictates the positioning of the yaw and pitch gyro combination and servounits. Similarly, roll stabilization is best assured by pivoting thewing flaps, or ailerons, 10b on the coplanar surfaces adjacent the rearend portion of the missile. This again dictates the position of the rollservo and gyro units, and of the tail antennae A and B mounted on theassociated cruciform surfaces, and serving as the range gating inputelements for feeding to the amplifiers 26 and 27 (FIG. 7) the timecontrol signal energy intercepted directly from the radar transmitter atthe launching vehicle or station; that is, the transmitter sendingoutthe radar pulses or a continuous wave, as the case might be, which arereflected back by the target to the phase comparison forward antennae AB A and B In this manner, this conventional range gate is able todetermine the target range. As will be noted in FIG. 7 the inputs 71 and72 from the tail antennas A and B to the range gate 70 are onlypartially illustrated but it is understood that any conventionalconnection can be used. The other major components then logically fallinto the relative locations indicated in FIG. 8. I

In lieu of supplying course maintaining signals to the control servos,the latter could be supplied with course changing signals from the sameseeker system, as on any occasion when collision avoidance, rather thancollision, should be desired. Also, in lieu of applying the invention tothe control of a pilotless missile, to steer it on a collision course,it could be applied to the control of a piloted aircraft, to steer itautomatically toward a landing point.

Obviously many modifications and variations of the present invention arepossible in the light of the above teachings. It is therefore to beunderstood that within the scope of the appended claims the inventionmay be practiced otherwise than as specifically described.

What is claimed is:

1. In a homing device, receiver means including a fixed antenna arraydisposed in a plane in orthogonal relation to the longitudinal axis ofsaid homing device and operative to retrieve a portion of a radiantenergy train transmitted to a distanct object and reflected back to saidantenna array by said object, and means for steering said homing deviceoperatively connected to said antenna array, said steering meansincluding a servo mechanism incorporating a power applying element whosemovement is proportional to the time rate of change in the angle ofincidence established by said reflected energy train as it impinges uponsaid antenna array.

2. In a homing device, receiver means including a pair of antennaelements in spaced relationship to the axis of the device, and to eachother, said antenna elements being operative to retrieve a portion ofthe energy transmitted to a distance object and reflected back by saidobject, and means for steering said homing device operatively connectedto said antenna array, said steering means incorporating a powerapplying element whose movement is proportional to the phase ratedifference between the energy received by one of said antenna elements,and that received by the other. I

3. Means for guiding a'missile to a maneuvering target, comprisingtwopairs of identical antennas mounted symmetrically about thelongitudinal axis of the missile, for reception of phase differentiatedsignals reflected from the target, and electronic means for converting,the phase difference between said signals into a measurement of thetime rate of change of the angle between the heading of the missile andthe line joining the target and the missile.

4. A system for directing a guided missile to intercept an illuminatedtarget, said system comprising a plurality of fixed antenna elementsspaced with respect to one another on the missile to receive signalreturn energy from said target, pairs of said'elements receiving signalsdiffering in phase in accordance with angular relationship of theinstantaneous missile axis path and the instantaneous target position,means responsive to said phase differing signals to provide a frequencydiscriminated signal representing the time rate of change of lead anglebetween the missile heading and the line of sight to the target andwhich determines the time'interv'al during which the.mis= sile should bemaintained in a circular path for first stage maneuvering, rate controlmeans responsive to said frequency discriminating and timing means tocause the missile to effect a second stage linear path of interceptionwith the target path of the second stage, the second stage beingcommenced when the angular velocity of the missile becomes equal to thetime rate of change of the missile angle,

5. A system for directing a guided missile to intercept a radarilluminated target, said system comprising a first pair of antennaeincluding a first and a second antenna and a second pair of antennaeincluding a third and fourth atenna, all four antennae being mountedsymmetrically on the sides of the missile, said antennae receivingreturn energy signals from the illuminated target and having a receivedsignal phase difference between said first and second antennae inaccordance with azimuth bearing of the target with respect to themissile, and having a received signal phase difference between saidthird. and fourth antennae in accordance with elevation bearing of Y thetarget with respect to the missile, a yaw and a pitch servo unit, meansto reduce said respective phase differing signals to an intermediatefrequency, first amplifier means to amplify said phase differingsignals, means to compare the phase of the first and second and of thethird and fourth antennae received signals to produce a phase relationdependent on the instantaneous phase relation of the signals from eachof the pairs of antennae, converting means to convert the output of saidcomparison means to a frequency modulated control signal in accordancewith the rate of phase variation of the signals at the antennae, meansto remove amplitude modulation of said converted signals, frequencydiscriminator means to detect the magnitude and sense of the phase rate,a gyroscope unit to introduce an axis deviation function, said gyroscopeunit comprising an integrating gyroscope whose precession is a measureof angular deviations of a missile airframe from the prescribedreference heading, a rotating signal generator, a second amplifier,angular velocity applied to the missile causing development of aprecession torque about the gyroscope axis, thereby rotating the signalgenerator to cause the generator to produce a voltage, said voltagebeing amplified in said second amplifier, a third amplifier, signalsfrom the frequency discriminator being added to the output of the secondamplifier to generate a servo command signal required for maintenance ofangular velocity dictated by the phase difference of incoming signals,and means responsive to said servo command signals to accordinglydeflect control surfaces of the missile.

6. Means to establish a collision course of a missile with a target,said means comprising spaced antennae means responsive to radiant energysignals phase differentiated in accordance with target to missile axisangular relation and received from the target, means to apply lateralacceleration to the missile in response to variations in the rate ofchange in the characteristics of intelligence reflected from thedetected target, phase difference between the signals varying as theangle between the target and the missile axis changes, the frequencydifference due to the changing angle being a measure of the time rate ofchange of the angle, to determine the time interval during which themissile should be maintained in the circular path constituting a firststage of maneuvering, a detection circuit, the two out-of-phase signalsreceived at pairs of the antennae being fed into said detection circuit,means to superimpose the output of said detection circuits upon acarrier wave, frequency modulation means, means to generate signalsrepresenting angular deviations of the missile airframe from prescribedreference heading and means responsive to frequency modulated signalsfrom said frequency modulation means to combine the frequency modulatedsignals with the signals representing angular deviations of the missileairframe from prescribed reference heading, adding of the two signalscausing the necessary servo command required for maintenance of angularvelocity to cause a target interception linear path.

References Cited UNITED STATES PATENTS 2,663,518 12/1953 Mulfiy 343-7 X2,644,397 7/1953 Katz 114-23 X 2,557,401 6/1951 Agins 24414 2,701,8752/1955 Baltzer 24414.2 2,420,016 5/ 1947 Sanders 244-77 2,826,380 3/1958Ketchledge 24414.3 2,751,494 6/1956 Gray 24414.3

OTHER REFERENCES Machover, Carl, Basics of Cyroscopes, John F. RiderPublisher, Inc., New York, 1960.

BENJAMIN A. BORCHELT, Primary Examiner US. Cl. X.R. 343-7

